The present invention relates to a system for ventilating or cooling the platforms of turbine blades.
Jet engine aircraft designers and manufacturers are continuously seeking ways to improve the performance of aircraft jet engines. One of the desired improvements is the lowering of specific fuel consumption of the jet engines, which may be accomplished by operating the engine at the highest possible temperatures at the turbine inlet. The specific fuel consumption may be increased by increasing the engine compression ratio which causes an increase in temperature at the outlet of the compressor. The increase in temperature at the turbine inlet also increases the engine thrust.
If these operating conditions can be achieved, it is possible to reduce the amount fuel carried by the aircraft to travel a certain distance, or to travel a longer distance given the same fuel quantity, and to lower the weight and size of the jet engine while the thrust remains the same.
However, the temperature at the turbine inlet is limited by the mechanical strength of the nozzle vanes and turbine blades. The mechanical properties of these elements are markedly reduced at extremely high temperatures. The rotor blades are also subjected to high centrifugal forces due to the high rotational speeds of the turbine rotor. In order to maintain the mechanical integrity of the turbine blades, the walls of the airfoil portion of the blades must be cooled, as must be the blade platforms which are attached to the blades and which form an inner boundary for the hot gas flowing over the airfoil portion of the blades.